US5394688A - Gas turbine combustor swirl vane arrangement - Google Patents

Gas turbine combustor swirl vane arrangement Download PDF

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Publication number
US5394688A
US5394688A US08/141,757 US14175793A US5394688A US 5394688 A US5394688 A US 5394688A US 14175793 A US14175793 A US 14175793A US 5394688 A US5394688 A US 5394688A
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United States
Prior art keywords
swirl
fuel
vanes
passage
swirl vanes
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US08/141,757
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English (en)
Inventor
David J. Amos
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Siemens Energy Inc
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Westinghouse Electric Corp
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Filing date
Publication date
Application filed by Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
Assigned to WESTINGHOUSE ELECTRIC CORPORATION reassignment WESTINGHOUSE ELECTRIC CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: AMOS, DAVID J., FOSS, DAVID T., LOWE, PERRY E., PARKER, DAVID M.
Priority to US08/141,757 priority Critical patent/US5394688A/en
Priority to TW083109273A priority patent/TW248585B/zh
Priority to US08/319,686 priority patent/US5479782A/en
Priority to AU75756/94A priority patent/AU7575694A/en
Priority to ES94307823T priority patent/ES2123102T3/es
Priority to DE69413352T priority patent/DE69413352T2/de
Priority to EP94307823A priority patent/EP0654639B1/en
Priority to CA002134419A priority patent/CA2134419A1/en
Priority to KR1019940027387A priority patent/KR950011818A/ko
Priority to CN94117610A priority patent/CN1107933A/zh
Priority to JP6289135A priority patent/JPH07180835A/ja
Publication of US5394688A publication Critical patent/US5394688A/en
Application granted granted Critical
Assigned to SIEMENS WESTINGHOUSE POWER CORPORATION reassignment SIEMENS WESTINGHOUSE POWER CORPORATION ASSIGNMENT NUNC PRO TUNC EFFECTIVE AUGUST 19, 1998 Assignors: CBS CORPORATION, FORMERLY KNOWN AS WESTINGHOUSE ELECTRIC CORPORATION
Assigned to SIEMENS POWER GENERATION, INC. reassignment SIEMENS POWER GENERATION, INC. CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: SIEMENS WESTINGHOUSE POWER CORPORATION
Assigned to SIEMENS ENERGY, INC. reassignment SIEMENS ENERGY, INC. CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: SIEMENS POWER GENERATION, INC.
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C1/00—Gas-turbine plants characterised by the use of hot gases or unheated pressurised gases, as the working fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C7/00—Combustion apparatus characterised by arrangements for air supply
    • F23C7/002—Combustion apparatus characterised by arrangements for air supply the air being submitted to a rotary or spinning motion
    • F23C7/004—Combustion apparatus characterised by arrangements for air supply the air being submitted to a rotary or spinning motion using vanes
    • F23C7/006—Combustion apparatus characterised by arrangements for air supply the air being submitted to a rotary or spinning motion using vanes adjustable
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/26—Controlling the air flow

Definitions

  • the present invention relates to a combustor for burning fuel in compressed air. More specifically, the present invention relates to a low NOx combustor for a gas turbine.
  • fuel is burned in compressed air, produced by a compressor, in one or more combustors.
  • combustors had a primary combustion zone in which an approximately stoichiometric mixture of fuel and air was formed and burned in a diffusion type combustion process. Additional air was introduced into the combustor downstream of the primary combustion zone.
  • the overall fuel/air ratio was considerably less than stoichiometric, the fuel/air mixture was readily ignited at start-up and good flame stability was achieved over a wide range in firing temperatures due to the locally richer nature of the fuel/air mixture in the primary combustion zone.
  • the inner liner enclosing the primary combustion zone is subject to over-heating and deterioration, especially at its outlet edge.
  • a gas turbine having a compressor section for producing compressed air and a combustion section in which the compressed air is heated.
  • the combustion section includes a combustor having (i) an air inlet in air flow communication with the compressor section, (ii) a plurality of first swirl vanes disposed in the air inlet for imparting a first swirl angle to at least a first portion of the compressed air, and (ii) first means for rotating each of the first swirl vanes into at least first and second positions, whereby the first swirl angle may be adjusted.
  • the air inlet comprises first and second passages and the first swirl vanes are disposed in the first passage.
  • the combustor further comprises a plurality of second swirl vanes disposed in the second passage for imparting a second swirl angle to a second portion of the compressed air and second means for rotating each of the second swirl vanes into at least first and second positions, so that the second swirl angle may be adjusted.
  • each of the first vanes is rotatable about a common axis with one of the second vanes.
  • FIG. 1 is a schematic diagram of a gas turbine employing the combustor of the current invention.
  • FIG. 2 is a longitudinal cross-section through the combustion section of the gas turbine shown in FIG. 1.
  • FIG. 3 is a longitudinal cross-section through the combustor shown in FIG. 2.
  • FIG. 4 is an isometric view of the air inlet portion of the combustor shown in FIG. 3, with the flow guide shown in phantom for clarity.
  • FIG. 5 is a transverse cross-section taken through lines V--V shown in FIG. 3.
  • FIG. 6 is a cross-section taken through line VI--VI shown in FIG. 5 and shows a portion of the combustor air inlet in the vicinity of the swirl vanes, except that in FIG. 6 the swirl vanes have been rotated from their position shown in FIG. 5 so as to be essentially oriented at 0° to the radial direction to allow viewing of the retainer pins in both vanes in a single cross-section.
  • FIG. 7 is a detailed view of the portion of FIG. 3 enclosed by the oval marked VII.
  • FIG. 8 is a cross-section taken through lines VIII--VIII shown in FIG. 6.
  • FIG. 9 is an alternate embodiment of the swirl vane support shown in FIG. 6.
  • FIG. 1 a schematic diagram of a gas turbine 1.
  • the gas turbine 1 is comprised of a compressor 2 that is driven by a turbine 6 via a shaft 26. Ambient air 12 is drawn into the compressor 2 and compressed.
  • the compressed air 8 produced by the compressor 2 is directed to a combustion system that includes one or more combustors 4 and a fuel nozzle 18 that introduces both gaseous fuel 16 and oil fuel 14 into the combustor.
  • the fuel is burned in the compressed air 8, thereby producing a hot compressed gas 20.
  • the hot compressed gas 20 produced by the combustor 4 is directed to the turbine 6 where it is expanded, thereby producing shaft horsepower for driving the compressor 2, as well as a load, such as an electric generator 22.
  • the expanded gas 24 produced by the turbine 6 is exhausted, either to the atmosphere directly or, in a combined cycle plant, to a heat recovery steam generator and then to atmosphere.
  • FIG. 2 shows the combustion section of the gas turbine 1.
  • a circumferential array of combustors 4, only one of which is shown in FIG. 4, are connected by cross-flame tubes 82, shown in FIG. 3, and enclosed by a shell 22.
  • Each combustor has a primary zone 30 and a secondary zone 32.
  • the hot gas 20 exiting from the secondary zone 32 is directed by a duct 5 to the turbine section 6.
  • the primary zone 30 of the combustor 4 is supported by a support plate 28.
  • the support plate 28 is attached to a cylinder 13 that extends from the shell 22 and encloses the primary zone 30.
  • the secondary zone 32 is supported by eight arms (not shown) extending from the cylinder 13. Separately supporting the primary and second zones 30 and 32, respectively, reduces thermal stresses due to differential thermal expansion.
  • a primary combustion zone 36 in which a lean mixture of fuel and air is burned, is located within the primary zone 30 of the combustor 4. Specifically, the primary combustion zone 36 is enclosed by a cylindrical inner liner 44 portion of the primary zone 30.
  • the inner liner 44 is encircled by a cylindrical middle liner 42 that is, in turn, encircled by a cylindrical outer liner 40.
  • the liners 40, 42 and 44 are concentrically arranged so that an inner annular passage 70 is formed between the inner and middle liners 44 and 42, respectively, and an outer annular passage 68 is formed between the middle and outer liners 42 and 44, respectively.
  • Cross-flame tubes 82 one of which is shown in FIG. 3, extend through the liners 40, 42 and 44 and connect the primary combustion zones 36 of adjacent combustors 4 to facilitate ignition.
  • a dual fuel nozzle 18 is centrally disposed within the primary zone 30.
  • the fuel nozzle 18 is comprised of a cylindrical outer sleeve 48, which forms an outer annular passage 56 with a cylindrical middle sleeve 49, and a cylindrical inner sleeve 51, which forms an inner annular passage 58 with the middle sleeve 49.
  • An oil fuel supply tube 60 is disposed within the inner sleeve 51 and supplies oil fuel 14 to an oil fuel spray nozzle 54.
  • the oil fuel 14 from the spray nozzle 54 enters the primary combustion zone 36 via an oil fuel discharge port 52 formed in the outer sleeve 48.
  • Gas fuel 16' flows through the outer annular passage 56 and is discharged into the primary combustion zone 36 via a plurality of gas fuel ports 50 formed in the outer sleeve 48.
  • cooling air 38 flows through the inner annular passage 58.
  • Compressed air from the compressor 2 is introduced into the primary combustion zone 36 by a primary air inlet formed in the front end of the primary zone 30.
  • the primary air inlet is formed by first and second passages 90 and 92 that divide the incoming air into two streams 8' and 8".
  • the first inlet passage 90 has an upstream radial portion and a downstream axial portion.
  • the upstream portion of the first passage 90 is formed between a radially extending circular flange 88 and the radially extending portion of a flow guide 46.
  • the downstream portion is formed between the flow guide 46 and the outer sleeve 48 of the fuel nozzle 18 and is encircled by the second inlet passage 92.
  • the second inlet passage 92 also has an upstream radial portion and a downstream axial portion.
  • the upstream portion of second passage 92 is formed between the radially extending portion of the flow guide 46 and a radially extending portion of the inner liner 44.
  • the downstream portion of second passage 92 is formed between the axial portion of the flow guide 46 and an axially extending portion of the inner liner 44 and is encircled by the upstream portion of the passage 92.
  • the upstream portion of the second inlet passage 92 is disposed axially downstream from the upstream portion of first inlet passage 90 and the downstream portion of second inlet passage 92 encircles the downstream portion of the first inlet passage 90.
  • a number of axially oriented, tubular primary fuel spray pegs 62 are distributed around the circumference of the primary air inlet so as to extend through the upstream portions of the both the first and second air inlet passages 90 and 92.
  • Two rows of gas fuel discharge ports 64 are distributed along the length of each of the primary fuel pegs 62 so as to direct gas fuel 16" into the air steams 8' and 8" flowing through the inlet air passages 90 and 92.
  • the gas fuel discharge ports 64 are oriented so as to discharge the gas fuel 16" circumferentially in the clockwise and counterclockwise directions.
  • a number of swirl vanes 84 and 86 are distributed around the circumference of the upstream portions of the air inlet passages 90 and 92.
  • a swirl vane is disposed between each of the primary fuel pegs 62.
  • the swirl vanes 84 in the inlet passage 90 impart a counterclockwise (when viewed in the direction of the axial flow) rotation to the air stream 8' so that the air forms a swirl angle B with the radial direction.
  • the swirl vanes 86 in the inlet passage 92 impart a clockwise rotation to the air stream 8" so that the air forms a swirl angle A with the radial direction.
  • the swirl imparted by the vanes 84 and 86 to the air streams 8' and 8" helps ensure good mixing between the gas fuel 16" and the air, thereby eliminating locally fuel rich mixtures and the associated high temperatures that increase NOx generation.
  • the outer annular passage 68 forms a secondary air inlet for the combustor through which air stream 8"' flows into the secondary zone 32.
  • a number of secondary gas fuel spray pegs 76 are circumferentially distributed around the secondary air inlet passage 68. According to an important aspect of the current invention, the secondary fuel pegs 76 are disposed within the secondary air inlet passage 68 and are radially oriented to ensure that all of the gas fuel 16"' is properly directed into the secondary air inlet passage.
  • the secondary fuel pegs 76 are supplied with fuel 16"' from a circumferentially extending manifold 74, shown best in FIG. 6.
  • Two rows of gas fuel discharge ports 78 are distributed along the length of each of the secondary fuel pegs 76 so as to direct gas fuel 16"' into the secondary air steams 8"' flowing through the secondary air inlet passage 68.
  • the gas fuel discharge ports 78 are oriented so as to discharge the gas fuel 16"' circumferentially in both the clockwise and counterclockwise directions. Because of the 180° turn made by the secondary air 8"' as it enters passage 68, the radial velocity distribution of the air will be non-linear. Hence, the spacing between the fuel discharge ports 78 may be adjusted to match the velocity distribution, thereby providing optimum mixing of the fuel and air.
  • a flame is initially established in the primary combustion zone 36 by the introduction of fuel, either oil 14 or gas 16', via the central fuel nozzle 18.
  • fuel either oil 14 or gas 16'
  • additional fuel is added by introducing gas fuel 16" via the primary fuel pegs 62. Since the primary fuel pegs 62 result in a much better distribution of the fuel within the air, they produce a leaner fuel/air mixture than the central nozzle 18 and hence lower NOx.
  • the fuel to the central nozzle 18 can be shut-off. Further demand for fuel flow beyond that supplied by the primary fuel pegs 62 can then be satisfied by supplying additional fuel 16"' via the secondary fuel pegs 76.
  • the swirl vanes 84 and 86 are oriented in opposition to each other so that the swirl angles A and B tend to cancel each other out, resulting in zero net swirl in the primary combustion zone 36.
  • the optimum angle for the swirl vanes 84 and 86 that will result in good mixing with a minimum of pressure drop will depend on the specific combustor design and is difficult to predict in advance. Therefore, according to an important aspect of the current invention, the swirl vanes 84 and 86 can be rotated into various angles.
  • the rotatability of the swirl vanes 84 and 86 is achieved by rotatably mounting the swirl vanes 84 and 86 in pairs along a common axis. In the preferred embodiment, this is accomplished by mounting alternate swirl vane pairs on shafts formed by the tubes 72 that supply fuel 16"' to the secondary fuel pegs 76--specifically, the fuel peg supply tubes 72 extend through close fitting holes 116 and 118 in the swirl vanes 84 and 86. The remaining swirl vane pairs are rotatably mounted on close fitting alignment bolts, such as the bolts 112 shown in FIG. 9, instead of on the secondary fuel peg supply tubes 72. In addition to allowing rotation of the swirl vanes, the alignment bolts 112 serve to clamp the assembly together and provide concentric alignment of flow guide 46 and the inner liner 44.
  • a pin 96 is installed in each swirl vane and extends into a hole 98 that is formed in either the flange 88, in the case of the swirl vanes 84, or in the radial portion of the flow guide 46, in the case of the swirl vanes 86.
  • the pins 96 lock the swirl vanes into a predetermined angular orientation.
  • a number of lock pin holes 98 are formed in the flange 88 for each swirl vane 84. These holes are arranged in an arc so that the angle of each swirl vane 84 can be individually adjusted by varying the hole into which the pin 96 is placed when the combustor is assembled.
  • a similar array of holes 98 are formed in the flow guide 46 to allow individual adjustment of the angle of the swirl vanes 86.
  • the angle of the swirl vanes 84 and 86 can be individually adjusted to obtain the optimum swirl angles A and B for the incoming air.
  • FIG. 9 shows an alternative embodiment of the current invention whereby all of the pairs of swirl vanes 84 and 86 are rotatably mounted on close fitting alignment bolts 112, instead of mounting alternating vane pairs on the secondary fuel peg supply tubes 72.
  • the head of each bolt 112 is secured to the flange 88 and a nut 114 is threaded onto the bolt to secure the assembly in place.
  • the fuel tubes 72 extend directly across the inlet of the passages 90 and 92 to the manifold 74.
  • the inner liner 44 Since the inner liner 44 is directly exposed to the hot combustion gas in the primary combustion zone 36, it is important to cool the liner, especially at its downstream end adjacent the outlet 71. According to the current invention, this is accomplished by forming a number of holes 94 in the radially extending portion of the inner liner 44, as shown in FIG. 3. These holes 94 allow a portion 66 of the compressed air 8 from the compressor section 2 to enter the annular passage 70 formed between the inner liner 44 and the middle liner 42.
  • an approximately cylindrical baffle 80 is located at the outlet of the passage 70 and extends between the inner liner 44 and the middle liner 42.
  • the baffle 80 is attached at its downstream end 108 to the downstream end of the middle liner 42 via spot welds 104.
  • the downstream end 108 of the baffle 80 could be attached to the middle liner 42 via a fillet weld.
  • the front end 106 of the baffle 80 is sprung loaded to bear against the outer surface of the inner liner 44.
  • the front end 106 of the baffle 80 extends upstream only about one-third the length of the inner liner 44. However, in some cases, it may be preferable to extend the front end 106 of the baffle 80 further upstream so that the baffle encircles the entire large diameter portion of the inner liner 44.
  • a number of holes 100 are distributed around the circumference of the baffle 80 and divide the cooling air 66 into a number of jets 102 that impinge on the outer surface of the inner liner 44.
  • the baffle 80 allows the cooling air 66 to be used much more effectively in terms of cooling the inner liner 44.
  • inwardly projecting snubber blocks 122 are distributed around the circumference of the baffle 80 to provide frictional damping for the inner liner 44, as shown in FIG. 7.
  • the snubbers 122 are preferably coated with a wear resistant coating.
  • the snubbers 122 are sized so that there is a clearance between them and the inner liner 44 at assembly.
  • the baffle 80 not only cools the inner liner 44 but reduces its vibration.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US08/141,757 1993-10-27 1993-10-27 Gas turbine combustor swirl vane arrangement Expired - Lifetime US5394688A (en)

Priority Applications (11)

Application Number Priority Date Filing Date Title
US08/141,757 US5394688A (en) 1993-10-27 1993-10-27 Gas turbine combustor swirl vane arrangement
TW083109273A TW248585B (cs) 1993-10-27 1994-10-06
US08/319,686 US5479782A (en) 1993-10-27 1994-10-07 Gas turbine combustor
AU75756/94A AU7575694A (en) 1993-10-27 1994-10-12 Gas turbine combustor
ES94307823T ES2123102T3 (es) 1993-10-27 1994-10-25 Alabes de torbellino para camara de combustion de turbina de gas.
DE69413352T DE69413352T2 (de) 1993-10-27 1994-10-25 Verstellbaren Wirbelleitschaufeln einer Brennkammer einer Gasturbine
EP94307823A EP0654639B1 (en) 1993-10-27 1994-10-25 Adjustable swirl vanes for combustor of gas turbine
KR1019940027387A KR950011818A (ko) 1993-10-27 1994-10-26 가스 터빈 연소기
CA002134419A CA2134419A1 (en) 1993-10-27 1994-10-26 Gas turbine combustor
CN94117610A CN1107933A (zh) 1993-10-27 1994-10-27 燃气轮机燃烧室
JP6289135A JPH07180835A (ja) 1993-10-27 1994-10-27 ガスタービン

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US08/141,757 US5394688A (en) 1993-10-27 1993-10-27 Gas turbine combustor swirl vane arrangement

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US08/319,686 Division US5479782A (en) 1993-10-27 1994-10-07 Gas turbine combustor

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US5394688A true US5394688A (en) 1995-03-07

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US08/141,757 Expired - Lifetime US5394688A (en) 1993-10-27 1993-10-27 Gas turbine combustor swirl vane arrangement
US08/319,686 Expired - Fee Related US5479782A (en) 1993-10-27 1994-10-07 Gas turbine combustor

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Application Number Title Priority Date Filing Date
US08/319,686 Expired - Fee Related US5479782A (en) 1993-10-27 1994-10-07 Gas turbine combustor

Country Status (10)

Country Link
US (2) US5394688A (cs)
EP (1) EP0654639B1 (cs)
JP (1) JPH07180835A (cs)
KR (1) KR950011818A (cs)
CN (1) CN1107933A (cs)
AU (1) AU7575694A (cs)
CA (1) CA2134419A1 (cs)
DE (1) DE69413352T2 (cs)
ES (1) ES2123102T3 (cs)
TW (1) TW248585B (cs)

Cited By (135)

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US5475979A (en) * 1993-12-16 1995-12-19 Rolls-Royce, Plc Gas turbine engine combustion chamber
EP0762057A1 (de) * 1995-09-01 1997-03-12 Mtu Motoren- Und Turbinen-Union MàœNchen Gmbh Einrichtung zur Vermischung von Brennstoff und Luft für Brennkammern von Gasturbinentriebwerken
WO1997017574A1 (en) * 1995-11-07 1997-05-15 Westinghouse Electric Corporation Gas turbine combustor with enhanced mixing fuel injectors
US5794449A (en) * 1995-06-05 1998-08-18 Allison Engine Company, Inc. Dry low emission combustor for gas turbine engines
WO1998049496A1 (en) 1997-04-30 1998-11-05 Siemens Westinghouse Power Corporation An apparatus for cooling a combuster, and a method of same
WO1999017057A1 (en) 1997-09-30 1999-04-08 Siemens Westinghouse Power Corporation ULTRA-LOW NOx COMBUSTOR
US5901555A (en) * 1996-02-05 1999-05-11 Mitsubishi Heavy Industries, Ltd. Gas turbine combustor having multiple burner groups and independently operable pilot fuel injection systems
US6047550A (en) * 1996-05-02 2000-04-11 General Electric Co. Premixing dry low NOx emissions combustor with lean direct injection of gas fuel
RU2160416C2 (ru) * 1999-12-01 2000-12-10 ООО "Самаратрансгаз" АО "ГАЗПРОМ" Камера сгорания газотурбинного двигателя
US6209325B1 (en) * 1996-03-29 2001-04-03 European Gas Turbines Limited Combustor for gas- or liquid-fueled turbine
US6405536B1 (en) * 2000-03-27 2002-06-18 Wu-Chi Ho Gas turbine combustor burning LBTU fuel gas
US6460344B1 (en) 1999-05-07 2002-10-08 Parker-Hannifin Corporation Fuel atomization method for turbine combustion engines having aerodynamic turning vanes
US20030014976A1 (en) * 2001-07-17 2003-01-23 Mitsubishi Heavy Industries Ltd. Pilot burner, premixing combustor, and gas turbine
RU2210026C1 (ru) * 2001-12-27 2003-08-10 Государственное унитарное предприятие Тушинское машиностроительное конструкторское бюро "Союз" - дочернее предприятие Федерального государственного унитарного предприятия Российской самолётостроительной корпорации "МиГ" Акустическая топливная форсунка
US20030196440A1 (en) * 1999-05-07 2003-10-23 Erlendur Steinthorsson Fuel nozzle for turbine combustion engines having aerodynamic turning vanes
RU2217663C1 (ru) * 2002-11-25 2003-11-27 Открытое акционерное общество "Научно-производственное объединение "Сатурн" Кольцевая камера сгорания газотурбинного двигателя
US6666029B2 (en) 2001-12-06 2003-12-23 Siemens Westinghouse Power Corporation Gas turbine pilot burner and method
US6691515B2 (en) 2002-03-12 2004-02-17 Rolls-Royce Corporation Dry low combustion system with means for eliminating combustion noise
RU2224953C2 (ru) * 1999-05-31 2004-02-27 Нуово Пиньоне Холдинг С.П.А. Инжектор для жидкого топлива, предназначенный для горелок в газовых турбинах
US20040053181A1 (en) * 2000-10-16 2004-03-18 Douglas Pennell Burner with progressive fuel injection
WO2004053395A1 (de) * 2002-12-11 2004-06-24 Alstom Technology Ltd Verfahren und vorrichtung zur verbrennung eines brennstoffs
US20050252217A1 (en) * 2004-05-11 2005-11-17 Chen Alexander G Nozzle
US20060156735A1 (en) * 2005-01-15 2006-07-20 Siemens Westinghouse Power Corporation Gas turbine combustor
US7137256B1 (en) * 2005-02-28 2006-11-21 Peter Stuttaford Method of operating a combustion system for increased turndown capability
RU2300052C1 (ru) * 2005-11-01 2007-05-27 Сергей Александрович Маяцкий Форсунка с кислородной подпиткой
US20070224562A1 (en) * 2006-03-23 2007-09-27 Hiromitsu Nagayoshi Burner for combustion chamber and combustion method
US20070220898A1 (en) * 2006-03-22 2007-09-27 General Electric Company Secondary fuel nozzle with improved fuel pegs and fuel dispersion method
US20080245901A1 (en) * 2006-09-26 2008-10-09 Fady Bishara Vibration damper
EP1985924A1 (en) * 2007-04-23 2008-10-29 Siemens Aktiengesellschaft Swirler
US20090084082A1 (en) * 2007-09-14 2009-04-02 Siemens Power Generation, Inc. Apparatus and Method for Controlling the Secondary Injection of Fuel
US20090320485A1 (en) * 2006-05-12 2009-12-31 Nigel Wilbraham Swirler for Use in a Burner of a Gas Turbine Engine
US7665309B2 (en) 2007-09-14 2010-02-23 Siemens Energy, Inc. Secondary fuel delivery system
US20100180599A1 (en) * 2009-01-21 2010-07-22 Thomas Stephen R Insertable Pre-Drilled Swirl Vane for Premixing Fuel Nozzle
US20100269507A1 (en) * 2009-04-23 2010-10-28 Abdul Rafey Khan Radial lean direct injection burner
US20110000671A1 (en) * 2008-03-28 2011-01-06 Frank Hershkowitz Low Emission Power Generation and Hydrocarbon Recovery Systems and Methods
US20110027728A1 (en) * 2008-04-01 2011-02-03 Vladimir Milosavljevic Size scaling of a burner
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AU7575694A (en) 1995-05-18
CN1107933A (zh) 1995-09-06
KR950011818A (ko) 1995-05-16
DE69413352D1 (de) 1998-10-22
DE69413352T2 (de) 1999-05-12
CA2134419A1 (en) 1995-04-28
EP0654639B1 (en) 1998-09-16
ES2123102T3 (es) 1999-01-01
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US5479782A (en) 1996-01-02
TW248585B (cs) 1995-06-01
JPH07180835A (ja) 1995-07-18

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